CHAPTER 6 ENVIRONMENTAL CONDITIONS 6.1 Summary This Chapter provides the natural environment at Xichang Satellite Launch Center (XSLC), the thermal environment during satellite processing, the thermal and mechanical environment (vibration, shock & noise) experienced during the launch vehicle flight and the electromagnetic environment during ground processing and the flight. 6.2 Environment at XSLC The environmental data for the XSLC has been monitored over a significant period of time. Table 6-1 and Table 6-2 provide the average monthly temperature and average monthly ground wind speed respectively. Table 6-1 Average Monthly Temperature in XSLC Month Highest ( C) Lowest ( C) Average ( C) January 7.9 4.5 5.9 February 10.4 5.0 8.0 March 14.5 9.7 11.7 April 17.5 13.1 15.0 May 20.2 15.6 17.7 June 21.1 17.7 19.1 July 21.3 19.3 20.0 August 21.3 18.5 19.8 September 19.3 16.2 17.2 October 16.4 13.2 14.1 November 12.3 8.4 10.0 December 8.9 4.6 6.5 Issue 2011 6-1
Table 6-2 Average Monthly Ground Wind Speed in XSLC Month Average Speed (m/s) Days (>13 m/s) January 2.2 0.5 February 2.3 1.1 March 2.3 2.5 April 2.0 1.6 May 1.5 0.6 June 1.0 0.4 July 1.1 0.2 August 1.2 0.1 September 0.9 0.2 October 1.1 0.1 November 1.4 0.0 December 1.7 0.2 The relative humidity at launch site: Maximum: Minimum: 90% in rainy season 42% in dry season 6.3 Satellite Processing Environment Following arrival at XSLC, the satellite is integrated, tested and fueled in SC Processing Building BS2 and BS3, and then, it is transported to the launch pad for combined operations. The processing environments for the satellite therefore fall into three phases: processing in BS2 and BS3, transportation to launch pad and final preparation on launch tower. The following environment is provided and maintained throughout the whole process: Temperature: 15 C ~ 25 C Relative humidity: 35% ~ 55% Cleanliness: 100,000 class Acoustic level: 90 db For the detailed information about satellite processing environment, please refer to the XSLC User s Manual (Issue 2009). 6.3.1 Environment in SC Processing Building BS2 and BS3 During the integration and checkout of satellite in BS2 and BS3, the environment inside the building is maintained by air-conditioning system. 6-2 Issue 2011
6.3.2 Environment during Transportation to Launch Pad 6.3.2.1 For Fairing Encapsulation-on-Pad The satellite is transferred to the launch pad in an environmentally controlled and sealed container, and the transfer takes approximately 30 minutes. The satellite transfer container is a sealed cylinder with a height of 8,070 mm and diameter of 4,120 mm with its wall made of aluminum sandwich for thermal insulation (See Chapter 8). Before transportation, the container is pressurized with dry nitrogen, followed by a thermal stabilization period to allow the satellite and container to stay at a steady temperature between 15 C and 25 C. The container is then sealed and moved out of BS3 for transport to the pad. When the container arrives at the launch pad, it is hoisted to the 8th floor of the Service Tower, where an environmentally controlled clean room has been established. The procedures for this operation are detailed in Chapter 8. After arrival at the launch pad clean room, the container goes through another stabilization period for the satellite thermal conditions to become adapted to those of the clean room. When the satellite and container have reached a thermally stabilized status, the container is opened and the satellite is moved out to mate to the launch vehicle. 6.3.2.2 For Fairing Encapsulation-in-BS3 The same environment conditions as Paragraph 6.3.2.1 are provided and maintained except that the fairing replaces the SC container. When the satellite encapsulated with fairing arrives at the launch pad, it is hoisted to the 8th floor of the Service Tower. The fairing is mated to the launch vehicle. The procedures for this operation are detailed in Chapter 8. 6.3.3 Air-conditioning inside Fairing The fairing air-conditioning system is connected following the mating of the satellite encapsulated in the fairing to the launch vehicle (Encapsulation-in-BS3) or final encapsulation of the satellite at the launch pad (Encapsulation-on-Pad). The air-conditioning system becomes operational immediately following its connection to the fairing and the configuration is shown in Figure 6-1. The air-conditioning system maintains the environmental conditions within the fairing as Issue 2011 6-3
LM-3A Series Launch Vehicle User s Manual shown in Table 6-3. Table 6-3 Air-conditioning parameters inside Fairing Temperature 13 C ~ 23 C Relative Humidity 35% ~ 55% Cleanliness 100,000 class Air Speed inside Fairing 2 m/s Noise inside Fairing 90 db Air Flow Rate 1500 ~ 3000 m3/hour (adjustable) The air-conditioning is shut off 45 minutes before launch and can be re-established within 40 minutes if the launch is aborted. The air-conditioning inlets are shown in Chapter 4. Figure 6-1 6-4 Pre-launch Air-condition System of LM-3A Series LVs Issue 2011
6.4 Electromagnetic Environment The satellite has to withstand the electromagnetic field generated by both the launch site RF facility and launch vehicle RF systems while mounted on the launch vehicle. These transmissions have been minimized to a very low level in the vicinity of the launch complex. This section details the frequencies and levels of the potentially interfering radiation. 6.4.1 Radio Equipment onboard LM-3A Series Launch Vehicles and Ground Test Equipment Characteristics of on-board radio equipment and ground test equipment are shown in Table 6-4. Table 6-4 Characteristics of on-board radio equipment and ground test equipment Frequency Power Antenna Equipment Sensitivity Polarization (MHz) (W) position Telemetry Transmitter 3 2200~2300 10 / linear VEB GNSS receiver 1500~1700 / -130 dbw RHCP VEB LAUNCH VEHICLE GROUND Telemetry Transmitter 2 Responder 1 Responder 2 2200~2300 5 / linear Rx 5860~5910 Tx 4210~4250 Rx and Tx 5580~5620 2-120 dbw linear 300 (peak) 0.8~1.0 µs 800 Hz <300 mw (average) -90 dbw linear Beacon 2730~2770 2 / linear Telemetry command Receiver Tester for Responder 1 Tester for Responder 2 Telemetry Command Transmitter 550~750 / -128 dbw / 5870~5910 0.5 / / 5570~5620 100 W (peak) / / 550~750 1 W / / Stage-2 Intertank Stage-3 Rear shell Stage-3 Rear shell Stage-2 Intertank Tracking & safety system ground test room at launch pad The locations of the onboard radio equipment are shown in Figure 6-2. Issue 2011 6-5
Figure 6-2 Locations of On-board Radio Equipment 6.4.2 RF Equipment and Radiation Strength at XSLC Working frequency: 5,577 ~ 5,617 MHz Antenna diameter: 4.2 m Impulse power: <1,500 kw Impulse width: 0.0008 ms Minimum pulse duration: 1.25 ms Average power: <1.2 kw 6.4.3 Launch Vehicle Electromagnetic Radiation and Susceptibility The electromagnetic radiation levels generated by the launch vehicle and the launch site equipment will not exceed those shown in Figure 6.3. The launch vehicle electromagnetic susceptibility is shown in Figure 6.4. The energy levels of launch vehicle electromagnetic radiation and susceptibility are measured at 1 m above VEB. The electromagnetic radiation 6-6 Issue 2011
of satellite shall not exceed the values shown in Figure 6-4. Frequency (MHz) Field Strength (db V/m) 0.01-0.05 80 0.05-3 90 3-300 70 300-550 80 550-750 103 750-2200 80 2200-2300 138 2300-2730 80 2730-2770 107 2770-4200 80 4200-4250 107 4250-5580 80 5580-5620 99 5620-6000 80 6000-6500 80 6500-13500 80 13500-15000 80 15000-80 Figure 6-3 Electromagnetic Radiation from LV and XSLC Issue 2011 6-7
Frequency (MHz) Field Strength (db V/m) 0.01-550 134 550-750 15 750-1000 134 1000-1500 140 1500-1700 10 1700-5580 140 5580-5910 35 5910-100000 140 Figure 6-4 Launch Vehicle Electromagnetic Radiation Susceptibility 6.4.4 EMC Analysis CALT and the customer shall jointly conduct an EMC analysis for the satellite, launch vehicle and launch site to verify their compatibility. The customer shall use the data provided in this User s Manual and shall provide the following information, as required in the ICD, to CALT for analysis: Satellite RF system configuration, characteristics, operating and radiating time periods, antenna position and its transmission direction, etc. Values and curves of the electric field of intentional and parasitic radiation generated by the satellite RF system at satellite separation plane The values and curves of the electromagnetic susceptibility of the satellite. 6-8 Issue 2011
6.5 Mechanical Environments 6.5.1 Steady State Acceleration The launch vehicle flight generates external forces on the satellite due to the engine thrust and aerodynamic forces. The typical maximum longitudinal steady state acceleration during the launch vehicle powered flight is shown in Table 6-5. It can be seen that the maximum longitudinal static acceleration occurs during the first stage flight. The maximum static acceleration will vary slightly with the satellite mass. Table 6-5 Typical Maximum Static Longitudinal Acceleration LM-3A LM-3B LM-3C During First Stage Flight +5.0 g N/A N/A During First Stage Booster Flight N/A +5.3 g +5.3 g During First Stage Core Flight N/A +3.6 g +3.6 g During Second Stage Flight +2.9 g +2.8 g +2.8 g During Third Stage first powered flight +1.6 g +1.2 g +1.2 g During Third Stage second powered flight +2.7 g +2.5 g +2.5 g Note: + means the direction of the acceleration coincides with the launch vehicle +X axis. 6.5.2 Vibration Environments 6.5.2.1 Sinusoidal Vibration Low level sinusoidal vibration is present during the flight, but the significant vibration events occur during engine ignition and engine shut-off, transonic flight and the stage separations. The sinusoidal vibration (zero-peak value) at the satellite separation plane is shown in Table 6-6 below. Table 6-6 Sinusoidal Vibration Levels Direction Frequency Range Amplitude & Acceleration Longitudinal Lateral 5-8 Hz 3.11 mm 8-100 Hz 0.8 g 5-8 Hz 2.33 mm 8-100 Hz 0.6 g Issue 2011 6-9
6.5.2.2 Random Vibration The maximum random vibration levels are primarily generated by the noise during the launch vehicle lift-off and transonic flight periods. The acoustic spectrum defined in Paragraph 6.5.3 covers excitations produced by random vibration. 6.5.3 Acoustic Noise The acoustic noise spectrum includes the engine noise and aerodynamic noise during flight, with the maximum acoustic noise being experienced by the satellite during the lift-off and the transonic flight phase. The values in Table 6-7 below are the maximum noise levels inside the fairing from the launch and transonic flight phase. Table 6-7 Acoustic Noise in the Fairing Central Frequency of Octave Bandwidth (Hz) Acoustic Pressure Level (db) 31.5 124 63 129 125 134 250 138 500 133 1000 129 2000 128 4000 127 8000 122 Total Acoustic Pressure Level 141.5 Note: 0 db corresponds to 2 10-5 Pa. 6.5.4 Shock Environment The maximum shock experienced by satellite occurs at SC/LV separation. The shock response spectrum (Q=10) at the separation plane is shown in Figure 6-5 for the C100 clampband and Figure 6-6 for C60 clampband. Additionally, CGWIC could also support the customer to procure low shock level separation system (clampband and the pyros) developed by European suppliers, if needed. 6-10 Issue 2011
Frequency (Hz) Shock Response Spectrum (Q=10) 100~1000 10.5 db/octave 1000~4000 4000 g Figure 6-5 Shock Response Spectrum at SC/LV Separation Plane - C100 Clampband Figure 6-6 Shock Response Spectrum at SC/LV Separation Plane - C60 Clampband Issue 2011 6-11
6.6 Thermal Environment This section covers the thermal environment during the launch vehicle flight up to separation, which is described separately for the two flight phases, i.e. before fairing jettison and after fairing jettison. 6.6.1 Thermal Environment prior to Fairing Jettison During the ascent, the net thermal flux density radiated by the fairing does not exceed 500 W/m 2 at any point. This does not include any thermal input from the satellite due to power dissipation by the satellite electronics. 6.6.2 Thermal Environment Post Fairing Jettison The free molecular heating flux upon fairing jettison is not greater than 1,135 W/m 2. This flux is calculated as the free molecular flow acting on a plane surface perpendicular to the velocity vector of the launch vehicle. The typical free molecular thermal flux after fairing jettisoning is shown in Figure 6-7. The thermal effects caused by Sun radiation, Earth infrared radiation and Earth albedo are also considered in the thermal analysis and will be determined through the ICD and Mission Review inputs. Figure 6-7 Free Molecular Heating Flux of LM-3A/3B/3BE/3C 6-12 Issue 2011
6.6.3 Launch Vehicle Generated Heat Flux The second stage retro-rockets burn for approximately 1.5 s and will generate a heat flux of < 300 W/m 2 at the satellite separation plane. The heat flux with the third-stage engines in operation will not exceed 350 W/m 2 at the satellite separation plane. 6.7 Pressure Environment During the ascent phase of launch vehicle flight, the fairing is vented in order to balance the pressure inside and outside the fairing and ensure there is no significant pressure difference. The maximum depressurization rate inside fairing during launch vehicle flight will not exceed 6.9 kpa/s. The typical design range for fairing internal pressure is shown in Figure 6-8. Figure 6-8 Fairing Internal Pressure 6.8 Contamination and Cleanliness Control Issue 2011 6-13
6.8.1 Contamination Control During the satellite processing at XSLC, the maximum organic non-volatile contamination on the satellite shall not exceed 2 mg/m 2 /week from opening of the satellite container in BS up to launch. The materials used in the clean room facilities and the parts of the launch vehicle in the fairing have been chosen to minimize outgassing and ensure that the total organic non-volatile contamination on the satellite does not exceed 4 mg/m 2 from the satellite initial encapsulation up to its separation from the launch vehicle. 6.8.2 Cleanliness Control The cleanliness of satellite processing facilities is maintained to class 100,000. Any hardware has to be subjected to a cleaning process followed by an inspection before it enters the clean rooms. All the materials used on the launch vehicle in the vicinity of the satellite are designed not to accumulate dust and selected to be non-shedding during ground processing. Prior to encapsulation of the satellite, the fairing, SC Container and Payload Adapter (PLA) are cleaned and inspected. 6-14 Issue 2011